EP0210249B1 - Generateur a gaz a turbine radiale a entree double - Google Patents

Generateur a gaz a turbine radiale a entree double Download PDF

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Publication number
EP0210249B1
EP0210249B1 EP86901145A EP86901145A EP0210249B1 EP 0210249 B1 EP0210249 B1 EP 0210249B1 EP 86901145 A EP86901145 A EP 86901145A EP 86901145 A EP86901145 A EP 86901145A EP 0210249 B1 EP0210249 B1 EP 0210249B1
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EP
European Patent Office
Prior art keywords
compressor
stage
turbine
gas generator
pressure ratio
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired
Application number
EP86901145A
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German (de)
English (en)
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EP0210249A1 (fr
Inventor
Rolf Jan Mowill
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Ulstein Turbine AS
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Ulstein Turbine AS
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Publication date
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Publication of EP0210249A1 publication Critical patent/EP0210249A1/fr
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • F02C3/10—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with another turbine driving an output shaft but not driving the compressor
    • F02C3/103—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with another turbine driving an output shaft but not driving the compressor the compressor being of the centrifugal type
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/50—Application for auxiliary power units (APU's)

Definitions

  • the invention relates to a high efficiency spool gas turbine gas generator according to the preamble of claim 1.
  • a gas turbine gas generator is disclosed in EP-A-0 161 562 which forms the preamble of claim 1 and which is post published prior art pursuant to article 54 (3) for all designated contracting states.
  • a further gas turbine gas generator is known from US-A-4 244 191.
  • the cycle according to the present invention utilizes a substantially higher pressure ratio than conventional gas turbine engines. Since pressure ratio is one of the important factors contributing to efficiency, the cycle according to the invention will provide higher thermal efficiencies than existing turbine engines.
  • the mechanical strength of the radial inflow turbine is sufficient to allow operation of each of the centrifugal compressor stages at near optimum level in terms of specific speed resulting in an optimized overall compressor component.
  • the radial inflow turbine working under these conditions surprisingly operates at near its own optimum specific speed, resulting in an optimized gas generator.
  • the resulting high peripheral speed of the radial inflow turbine results in a lowering of the stagnation temperature of the hot combustion gases impinging on the turbine blades.
  • the turbine inlet temperatures can be raised to further increase thermal efficiency or the turbine component operating lifetime can be extended.
  • the gas generator overall pressure ratio across the first and second compressor stages is greater than about 20 : 1, and the first compressor stage pressure ratio is between about 6 : 1 and about 9 : 1, while the second compressor stage pressure ratio is between about 2 : and about 4 : 1.
  • each compressor stage ranges from about 0.65 - 0.85.
  • Fig. 1 is a schematic cross-sectional view of a gas turbine gas generator in accordance with the present invention.
  • Figs. 2A and 2B are schematic graphs showing the improvements in efficiency resulting from use of compressor components and turbine components matched to their specific speeds in a gas generator made in accordance with the present invention.
  • the gas generator of the present invention includes compressor means for providing an overall pressure ratio of greater than about 15 : 1 and having a double-entry centrifugal low pressure first stage with a pair of entrances and a single exit.
  • compressor means 12 including low pressure, first compressor stage 14.
  • First compressor stage 14 includes double-entry compressor module 16 having a compressor rotor 18 mounted for rotation by shaft assembly 20 and with twin, axially-opposed flow paths designated by arrows 22, 24, respectively.
  • First compressor stage 14 also includes surrounding housing 26 which defines a fair of flow-symmetric, axially opposed entrances 28, 30 for directing air to compressor blade assemblies 32, 34 on rotor 18.
  • An improved double-entry compressor especially suited for this application is disclosed in US patent no 4 530 639.
  • Double-entry compressor module 16 has a single, annular, radially directed compressor exit 36 operatively connected to a diffuser assembly 38.
  • Diffuser assembly 38 receives the high velocity air from exit 36 and converts the high velocity air to higher pressure, low velocity air for ultimate transmission to high pressure compressor stage 40 and then to combustor 60.
  • Diffuser assembly 38 could be replaced by a manifold assembly (not shown) designed to preserve a portion of the dynamic head of the air leaving exit 36.
  • a variable geometry diffuser apparatus which advantageously could be used in diffuser assembly 38 is disclosed in US patent no 4 573 868.
  • diffuser assembly 38 is provided with an annular plenum 42 for collecting the diffused air.
  • plenum 42 is flow-connected with cross-over duct 44 which, in turn, is flow-connected with the inlet plenum 46 of high pressure compressor stage 40 to be discussed in more detail hereinafter.
  • cross-over duct arrangements can be utilized including the construction detailed in EP-A-0 161 562, which construction offers additional structural and component performance benefits.
  • the compressor means includes a high pressure centrifugal compressor second stage positioned adjacent to the first stage compressor and having a single entrance and a single exit.
  • compressor means 12 further includes high pressure compressor stage 40 which is a single-entry radial compressor having a housing 48 defining compressor entrance 50 for receiving compressed air from inlet plenum 46.
  • the high pressure compressor housing 48 also defines second stage compressor exit 52 in flow communication with the combustor supply plenum 54 through second stage diffuser assembly 56.
  • High pressure compressor rotor 58 is positioned within housing 48 and is mounted for rotation on shaft assembly 20.
  • the compressed air flow path proceeds from the collection plenum 42 of the double-entry compressor module 16, through cross-over duct 44, to the high pressure compressor stage inlet plenum 46, past high pressure compressor rotor 58 and into combustor supply plenum 54.
  • the pressure ratios of the first and second compressor stages are selected such that the pressure ratio in the first, low pressure stage, is greater than about twice the pressure ratio of the second, high pressure stage. Additionally, the flow path dimensions of the first compressor stage are selected to provide a favorable specific speed consistent with the pressure ratio chosen for that stage, as will be explained in more detail hereinafter. As embodied herein, the overall compression ratio across both first and second compressor stages is greater than about 15 : 1, with the pressure ratio of first compressor stage 14 ranging from about 6 : 1 to about 9 : 1 and second compressor stage 40 pressure ratio ranging from about 2 : 1 to about 4 : 1. The relatively low second stage pressure ratio is intended to keep the specific speed (see discussion below) of the second stage as high as possible.
  • the typical inducer tip relative entrance Mach numbers at first stage compressor entrances 28, 30 are about 1.4 or greater and occur at the outer tips of the leading edges of blades 32, 34 during operation at rated power. Although greater than 1.0 Mach numbers will cause shocks to occur in the compressor inlet, these will be relatively weak, oblique shocks and will not seriously affect overall performance.
  • the present invention stems from a design philosophy where control of specific speed of the compressor components is more important than keeping low inlet Mach numbers. This is a departure from conventional design practice in the art.
  • N s The specific speed (N s ) is here defined as follows:
  • the gas generator includes combustor means operatively connected to the exit of the second compressor stage for receiving the compressed air and combusting fuel using the compressed air to generate combustion gases.
  • combustor means operatively connected to the exit of the second compressor stage for receiving the compressed air and combusting fuel using the compressed air to generate combustion gases.
  • two can combustors 60 are provided in flow-communication with plenum 54 through double-walled turbine inlet manifold 62 (only a single combustor is shown in Fig. 1). Any number of combustors can be used and also combustors of non-circular cross-section.
  • all the compressed air exiting second compressor stage 40 is channeled to combustor 60 via plenum 54 and between the walls of manifold 62.
  • the gas generator includes a single stage radial inflow turbine having an inlet and an outlet.
  • the turbine is operatively connected to directly drive the shaft assembly on which the compressor stages are mounted and is also flow connected to the combustor means for receiving at the turbine inlet, and partially expanding, the combustion gases.
  • gas generator 10 includes radial inflow turbine 66 with turbine rotor 68 having blades 70 with blade tips 70a.
  • Turbine 66 receives the hot combustion gases from combustor 60 through turbine inlet nozzle assembly 72 via turbine inlet manifold 62.
  • Turbine rotor 68 is directly coupled to shaft assembly 20 to rotate both rotor 58 of high pressure compressor stage 40 and rotor 18 of low pressure compressor stage 14.
  • EP-A-0 161 562 provides details of apparatus especially suited for shaft assembly 20.
  • the gas generator includes exhaust means flow connected to the turbine outlet for ducting the partially expanded combustion gases for further external work-producing expansion.
  • gas generator 10 includes manifold 74 operatively connected to receive the partially expanded combustion gases from turbine 66 for ducting, e.g., to downstream free power turbine unit 90 or free-jet propulsion apparatus 92 (Fig. 1A), to provide further work-producing expansion of the combustion gases exhausted from turbine 66.
  • Free power turbine unit 90 can have one or more individual stages (two stages are shown in Fig. 1).
  • Fig. 2 the specific speed curves for the two compressor stages and the radial turbine stage of a gas generator designed according to the invention are shown.
  • the gas generator has an overall pressure ratio of greater than 20 : 1, using a double-entry centrifugal compressor of about 9: 1 pressure ratio with inlet Mach number greater than 1.4, and a second stage, single-entry centrifugal compressor of about 3.5 : 1 pressure ratio (i.e. less than 50 % of the pressure ratio of the first stage).
  • the efficiency of the radial inflow turbine peaks near the efficiency peak of the two compressors stages when all are run together on the same shaft as part of a gas generator. This is denoted with regions designated "A" in Fig. 2 which encompass compressor specific speed ranges from about 0.65 - 0.85 and a radial inflow turbine specific speed range of about 0.50 - 0.75.
  • Regions "B” indicate that a second stage centrifugal compressor following one or more axial stages or a single-entry centrifugal compressor, as hypothetical alternative low pressure "first stages” made in accordance with the teachings of prior art high pressure ratio gas generators, would fall well below its optimum range.
  • Fig. 2 also shows that a radial inflow gas generator turbine would fall below its optimum if it had to be matched to a lower specific speed compressor section. Additionally, because all the rotating components of the gas generator flow path according to the invention run at high and optimum specific speeds, the physical dimensions of the rotors become small and inexpensive to produce. Also, the inertia of the rotor assembly becomes small, which will facilitate quick starts.
  • the double-entry, first stage compressor will provide twice the mass flow compared with a single-entry compressor of comparable dimensions. Under steady state operating conditions when the air is compressed in the first stage to design levels, the flow path geometry of the latter will accommodate the entire first stage compressed gas flow rate. During start-up, however, and before the first stage volume flow is reduced to the design level, the second stage compressor cannot "swallow" the larger-than- designed-for volume flow.
  • bleed duct 76 is connected to cross-over duct 44 to receive compressed air from first compressor stage 14 when valve 78 is activated during start-up.
  • Automatic controller 80 is shown controlling valve 78, but manual operation could be substituted for controller 80.
  • controller 80 One skilled in the art could readily select appropriate apparatus for providing the function of controller 80 given the present disclosure.
  • the bled air can simply be discharged to the atmosphere as shown in Fig. 1.
  • the stagnation temperature of the combustion gases impinging on the turbine blade tips 70a, and hence the metal temperature of the turbine is much lower (about 100 - 200° C iower) than in an axial turbine stage exposed to the same nozzle inlet temperature.
  • the high blade tip speeds of the radial inflow turbine in the present invention which are dictated by the compressor requirements, actually provide an additional benefit in terms of lower metal temperatures. This enables higher turbine inlet temperatures to be used which, together with the high pressure ratio of the present invention, result in a very low specific fuel consumption.
  • the high pressure ratio cycle of the present invention can provide high thermal efficiency at firing temperatures up to about 1200° C with current materials without introducing cooling in the turbine rotor. At this temperature level all current axial turbines need cooling.
  • gas generators according to this invention utilize pressure ratios well above 20 : 1, it could be desirable to cool the radial turbine or to use non-metallic materials for the rotor, since the efficiency would be further improved. Cooling could also be applied in order to increase specific power even if thermal efficiency would not be improved.
  • the gas generator depicted in Fig. 1 would ideally operate with an overall pressure ratio of about 20 : 1, produce an equivalent shaft power of about 500 kw with about a 1.8 kg/s (4.0 Ib/s) air flow and turbine inlet temperature of greater than about 1200° C (2200° F); have a thermal efficiency of about 35 %; and use a gas generator turbine expansion ratio of about 5.0 : 1. Turbine rotor speed would be approximately 92,000 rpm equivalent to a turbine blade tip speed of about 750 m/s.
  • the projected equivalent engine fuel consumption would be about 0.21 - 0.24 kg/kwh (0.35 - 0.40 lb/hp/h), which is comparable to gas generators in state-of-the-art recuperated gas turbine engines under development.
  • the specific weight of engines using the present gas generator is only about 10 % that of a comparable diesel engine.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

Un générateur à turbine à gaz monorotor de rendement élevé utilise un compresseur centrifuge à double entrée dans un premier étage, un compresseur centrifuge à entrée unique dans un deuxième étage, et une turbine à écoulement radial pour obtenir un rapport de pression global supérieur à 15:1. Le rapport de pression dans le premier étage est supérieur au double de la valeur du rapport de pression dans le deuxième étage, et le nombre de Mach à l'entrée du premier étage est supérieur à 1,4 environ. La vitesse spécifique de chaque compreseur est comprise entre 0,65 et 0,85 et celle de la turbine entre 0,50 et 0,75.

Claims (9)

1. Générateur à gaz à turbine à gaz à simple enroulement à grand rendement qui comprend:
a) un dispositif de compression (12) comprenant:
(1) un compresseur d'air par centrifugation (14) à double entrée situé dans un premier étage et comprenant une paire d'entrées (28, 30) et une sortie commune (36),
(2) un compresseur d'air par centrifugation (40) situé dans un deuxième étage à proximité immédiate du compresseur (14) du premier étage, l'étage contenant le deuxième compresseur comportant une entrée reliée, pour écoulement, à la sortie commune du premier étage et également une sortie pour le deuxième étage, et
(3) un arbre (20) reliant mécaniquement le premier et le deuxième étages pour assurer leur rotation à la même vitesse angulaire.
b) Un dispositif de combustion (60) relié, pour le fonctionnement, à la sortie du deuxième étage pour recevoir l'air comprimé et assurer la combustion du combustible au moyen de l'air comprimé pour produire les gaz de combustion.
c) Une turbine radiale (66) à un seul étage comportant une entrée et une sortie, cette turbine étant reliée directement, pour le fonctionnement, à l'arbre d'entraînement (20) et étant également reliée, pour l'écoulement, au dispositif de combustion (60) pour recevoir à l'entrée de la turbine et détendre partiellement les gaz de combustion, et
d) un dispositif d'évacuation relié, pour l'écou- tement, à la sortie de la turbine pour diriger les gaz de combustion partiellement détendus vers une autre détente extérieure produisant du travail,
caractérisé en ce que:
(1) le taux de compression dans le premier étage de compression (14) est environ le double du taux de compression dans le deuxième étage de compression (40) et que le taux de compression d'ensemble est supérieur à 15 : 1,
(2) que les vitesses spécifiques dans le premier et dans le deuxième étages de compression (14, 40) sont supérieures chacune à 0,60, et
(3) que la vitesse spécifique de la turbine radiale (66) est comprise entre environ 0,50 et environ 0,75.
2. Générateur à gaz selon la revendication 1, caractérisé en ce que la vitesse spécifique dans chacun des étages de compression, le premier et le second (14, 40) tombent entre environ 0,65 et environ 0,85.
3. Générateur à gaz selon l'une des revendications 1 ou 2, caractérisé en ce que le taux de compression dans l'ensemble du premier et du deuxième étages de compression (14, 40) est supérieur à 20 : 1 environ.
4. Générateur à gaz selon l'une des revendications 1 à 3, caractérisé en ce que le taux de compression dans le premier étage de compression (14) est compris entre environ 6 : 1 et environ 9 : 1.
5. Générateur à gaz selon l'une des revendications 1 à 4, caractérisé en ce que le taux de compression dans le deuxième étage de compression (40) est compris entre environ 2 : 1 et environ 4 : 1.
6. Générateur à gaz selon l'une des revendications 1 à 5, caractérisé en ce que pratiquement la totalité de l'air comprimé qui sort du deuxième étage de compression (40) arrive dans le dispositif de combustion (60).
7. Générateur à gaz selon l'une des revendications 1 à 6, caractérisé en ce que le premier et le deuxième étages de compression (14, 40) sont disposés coaxialement par rapport à la turbine radiale (66) et en ce que cette turbine tourne à une vitesse périphérique supérieure à 700 m/s environ.
8. Générateur à gaz selon l'une des revendications 1 à 7, caractérisé en ce que le nombre de Mach à l'entrée de l'extrémité de l'inducteur dans le compresseur à double entrée (14) situé au premier étage est d'environ au moins 1 : 4 ou supérieur.
9. Générateur à gaz selon l'une des revendications 1 à 8, caractérisé en ce qu'il comporte un dispositif d'évacuation d'une partie de l'air comprimé sortant du premier étage de compression pendant le démarrage.
EP86901145A 1985-02-05 1986-02-05 Generateur a gaz a turbine radiale a entree double Expired EP0210249B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US698586 1985-02-05
US06/698,586 US4641495A (en) 1985-02-05 1985-02-05 Dual entry radial turbine gas generator

Publications (2)

Publication Number Publication Date
EP0210249A1 EP0210249A1 (fr) 1987-02-04
EP0210249B1 true EP0210249B1 (fr) 1989-11-15

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US (1) US4641495A (fr)
EP (1) EP0210249B1 (fr)
CA (1) CA1262638A (fr)
DE (1) DE3666966D1 (fr)

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US5081832A (en) * 1990-03-05 1992-01-21 Rolf Jan Mowill High efficiency, twin spool, radial-high pressure, gas turbine engine
US5377483A (en) * 1993-07-07 1995-01-03 Mowill; R. Jan Process for single stage premixed constant fuel/air ratio combustion
US5628182A (en) * 1993-07-07 1997-05-13 Mowill; R. Jan Star combustor with dilution ports in can portions
US5638674A (en) * 1993-07-07 1997-06-17 Mowill; R. Jan Convectively cooled, single stage, fully premixed controllable fuel/air combustor with tangential admission
US5613357A (en) * 1993-07-07 1997-03-25 Mowill; R. Jan Star-shaped single stage low emission combustor system
US5572862A (en) * 1993-07-07 1996-11-12 Mowill Rolf Jan Convectively cooled, single stage, fully premixed fuel/air combustor for gas turbine engine modules
US6220034B1 (en) 1993-07-07 2001-04-24 R. Jan Mowill Convectively cooled, single stage, fully premixed controllable fuel/air combustor
US5924276A (en) * 1996-07-17 1999-07-20 Mowill; R. Jan Premixer with dilution air bypass valve assembly
US6925809B2 (en) 1999-02-26 2005-08-09 R. Jan Mowill Gas turbine engine fuel/air premixers with variable geometry exit and method for controlling exit velocities
US6460324B1 (en) * 1999-10-12 2002-10-08 Alm Development, Inc. Gas turbine engine
US6397576B1 (en) * 1999-10-12 2002-06-04 Alm Development, Inc. Gas turbine engine with exhaust compressor having outlet tap control
US6363708B1 (en) * 1999-10-12 2002-04-02 Alm Development, Inc. Gas turbine engine
US7003961B2 (en) * 2001-07-23 2006-02-28 Ramgen Power Systems, Inc. Trapped vortex combustor
US7603841B2 (en) * 2001-07-23 2009-10-20 Ramgen Power Systems, Llc Vortex combustor for low NOx emissions when burning lean premixed high hydrogen content fuel
US6694743B2 (en) 2001-07-23 2004-02-24 Ramgen Power Systems, Inc. Rotary ramjet engine with flameholder extending to running clearance at engine casing interior wall
US7104068B2 (en) * 2003-08-28 2006-09-12 Siemens Power Generation, Inc. Turbine component with enhanced stagnation prevention and corner heat distribution
US6968697B2 (en) * 2003-09-17 2005-11-29 Honeywell International Inc. Integral compressor housing of gas turbine engines
US7628018B2 (en) * 2008-03-12 2009-12-08 Mowill R Jan Single stage dual-entry centriafugal compressor, radial turbine gas generator
AU2009352301B2 (en) 2009-09-13 2015-07-30 Lean Flame, Inc. Inlet premixer for combustion apparatus
US9651138B2 (en) 2011-09-30 2017-05-16 Mtd Products Inc. Speed control assembly for a self-propelled walk-behind lawn mower
US9033670B2 (en) 2012-04-11 2015-05-19 Honeywell International Inc. Axially-split radial turbines and methods for the manufacture thereof
US9115586B2 (en) 2012-04-19 2015-08-25 Honeywell International Inc. Axially-split radial turbine
EP2943668B1 (fr) * 2013-01-10 2018-04-04 United Technologies Corporation Générateur de gaz pourvu d'une armature ayant des passages d'air
US9476305B2 (en) 2013-05-13 2016-10-25 Honeywell International Inc. Impingement-cooled turbine rotor
US10202856B2 (en) * 2014-09-02 2019-02-12 United Technologies Corporation Decoupled gas turbine engine
US11788464B2 (en) 2019-05-30 2023-10-17 Joseph Michael Teets Advanced 2-spool turboprop engine
GB201915309D0 (en) * 2019-10-23 2019-12-04 Rolls Royce Plc Turboshaft

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Also Published As

Publication number Publication date
US4641495A (en) 1987-02-10
CA1262638A (fr) 1989-11-07
EP0210249A1 (fr) 1987-02-04
DE3666966D1 (de) 1989-12-21

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